This invention relates to gas turbine engines and, more particularly, to an inlet for use on a gas turbine engine whereby the operational noise level can be reduced during aircraft take-off and approach operation.
It is well understood that significant sources of noise generated by gas turbine engines are the result of viscous wake and rotor turbulence interaction. And, it is well known that the noise generated thereby is affected by such parameters as blade rotational speed, blade-to-blade spacing, blade geometry and by the proximity of stationary hardware to rotating blade rows. For example, the viscous wake interaction between the wakes of stationary upstream blade rows and moving downstream rows (rotors) results in noise at the rotor blade passing frequency and its harmonics which propagates at the speed of sound with respect to the fluid medium.
There are two currently popular approaches toward suppressing noise generated in this manner. One approach is to line the engine inlet with large amounts of sound-absorbing paneling. This technique is now well known in the art of acoustics, one such scheme being fully disclosed in U.S. Pat. No. 3,542,152-- Adamson et al, which is assigned to the assignee of the present invention. However, because of the close proximity of the fan or compressor to the inlet frontal plane, and the lack of acoustic shielding in the forward direction, a significant portion of the noise may still propagate forward out of the inlet duct. Generally, in order to obtain adequate noise suppression, more duct area must be lined than is practically available on the surface of the inlet walls. This has resulted in designs which incorporate acoustically treated, annular flow splitter rings in the inlet flow path with potential performance and foreign object ingestion problems.
Another approach, and one more related to the subject matter of the present invention, is to contour the inlet walls to accelerate the flows to sonic or near sonic velocity. The principle employed is that an acoustic wave cannot propagate upstream against sonic flow since the wave itself can only travel at sonic velocity. This, however, presents some performance problems since a considerable amount of inlet area variation is required because of the large variation in airflow with engine power setting. In particular, the problem is that of reducing the inlet throat area during partial power approach conditions while providing adequate throat area at the higher power take-off and cruise conditions. Several configurations have been proposed. In some schemes, the inlet is sized and contoured for cruise operation, but is provided with means for obdurating the inlet to increase the flow velocity at take-off and approach operation. One such scheme is fully disclosed in U.S. Pat. No. 3,611,724-- J. T. Kutney, assigned to the assignee of the present invention, and wherein an inflatable diaphragm is provided at the inlet throat to vary the cross-sectional area of a function of engine operational mode. However, these previous design configurations either have not been as aerodynamically clean as is desired or they have been mechanically complicated. It will therefore be appreciated that although the concept of an accelerating inlet (i.e., one in which the flow is accelerated to sonic or near-sonic velocity) is not new per se, a need exits for applying the concept to a gas turbine engine in an efficient, practical and viable manner.